1. Field of the Invention
The present invention relates to minimizing vibrations or oscillations of a body moving through a fluid at transonic speeds due to shock wave instabilities. In particular, the present invention relates an apparatus and method of predictably forming and fixing the location era shock wave on a body moving through the air such as a wing mounted pod on an airborne aircraft, thereby minimizing induced vibration.
2. Description of the Related Art
Vibrations resulting from unsteady transonic flow is a serious problem for aircraft traveling at subsonic/transonic velocities. Induced vibrations are particularly acute when associated with relatively thick bodies such as instrument pods and weapons systems suspended from wings on many military aircraft. Transonic induced vibration levels can become so severe that they can pose a threat to the safety of the aircraft and crew. The vibrations encountered at transonic flow velocities are largely attributable to unstable interaction between the shock wave and the turbulent flow region trailing behind the shock wave. The flow behind the shock wave no longer follows the contours of the body and is called a separated flow or separated boundary layer. The fluid in the separated flow region and in the wake of the body is turbulent. To better understand the source of the vibrations and their relationship to the shock wave/separated flow interactions, a brief review of the progression from subsonic through transonic flow is described.
Consider the case of a smooth body 10 of FIG. 1 with a leading edge 12 pointing directly into the flow and a trailing edge 14 traveling through a semi-viscous medium such as air with a flow direction indicated by the arrow labeled 16. FIGS. 1A through 1D illustrate four stages of local transonic flow that such a smooth body 10 will experience as the freestream flow velocity is gradually increased. In the sequence shown in FIG. 1A trough 1D the freestream velocities range from values at which sonic conditions are first achieved near the body surface to freestream velocities at the upper end of transonic flow just below the supersonic range. The freestream region 18 of the flow is the flow that exists in the absence of influence from the body. The velocity of the flow in the freestream region is called the freestream velocity. The term semi-viscous is used because even largely inviscid fluids such as air exhibit viscosity when flowing around a body due to the shear stress that exist in the boundary layer 20 of the body 10. The boundary layer 20 illustrated in FIG. 2 is a layer of fluid in direct contact with the body surface 24 and extend some distance above the surface. In this layer 20, the velocity of the fluid flow increases rapidly as a function of distance from the body 10 as indicated by the length of the arrows collectively designated 22. The flow at the body surface 24 is zero. Above the upper limit of the boundary layer 26, the flow velocity is no longer a strong function of the distance from the body 10. Above the upper limit 26 the flow velocity as indicated by the arrow 28 corresponds to the local flow velocity in the vicinity of the body 10.
Fluid flow around smooth bodies at subsonic and low transonic freestream flow velocities in both inviscid and viscous fluids is ideally laminar as illustrated in cross-section in FIG. 1A. The freestream region 18 of the flow encounters the leading edge 12 of the body 10 and transitions to a flow around the body 10 that is largely laminar. Ideally, the flow around the body is smooth and not separated following along the body 10 contour. In reality, a vortex wake 19 with some flow separation and minor turbulence is almost always present at the trailing edge 14 of the body 10 even at very low flow velocities. The presence of the wake 19 does not generally change the nature of the smooth, non-separated flow over much of the rest of the body 10. While smooth, the flow velocity around the body is not constant due to the requirement of conservation of mass. Velocities of the flow near the body 10 increase from the freestream value at the forward or leading edge 12 to a maximum velocity near the thickest point, hereinafter called the high point 30 of the body 10. In the ideal situation, flow velocity then decreases until the trailing edge 14 is encountered and the freestream velocity is reestablished. Under real, non-ideal flow conditions the smooth decrease in flow velocity is interrupted by the vortex wake 19 somewhere before the flow reaches the trailing edge 14.
As the freestream velocity increases from purely subsonic to transonic velocities, a critical velocity M.sub.crit is eventually reached where the local flow near high point 30 reaches the speed of sound, or 1 Mach, and becomes sonic. This condition is illustrated in FIG. 1A and is characterized by the formation of a sonic point 32 at the high point 30. The sonic point 32 is a small region or point of flow that has reached the speed of sound. The flow elsewhere on the body is not appreciably altered by sonic point 32 formation. Generally, bodies have a pair of the high points 30 on opposite sides of the body 10. A cylinder or other linearly extended bodies, has a pair of high points that are actually high regions with finite lengths orientated perpendicularly with respect to the flow direction. A sphere has a high point ring encircling the sphere.
A further increase in the freestream velocity extends the region of local supersonic flow downstream as illustrated in FIG. 1B and results in the formation of shock waves 34 and 35 at a trailing edge of the sonic point 32 which has expanded to become a sonic flow region 36. The shock waves 34 and 35 begin to form when the local flow velocity exceeds about 1.2 Mach. The leading edge boundary of the supersonic flow region 36 is called the sonic line 37 for both of the shock waves 34 and 35 in FIG. 1B. Due to the viscous nature of the fluid flow near the body 10, the shock wave/boundary layer interaction initially thickens the boundary layer 20 and then results in the separation of the flow from the surface 24. Initial onset of separation begins with a characteristic separation bubble 38 forming just behind the shock waves 34 and 35. The wake vortex region 19 is typically widened by the separation bubble 38 formation. In some cases at velocities near that required for initial shock wave 34, 35 formation, the flow may reattach downstream from the separation bubble 38.
FIGS. 1C and 1D illustrate what happens when the freestream velocity has increased enough to result in massive boundary layer flow separation 40 behind the shock wave. At this stage, there is little chance for flow re-attachment to the body 10. The separated flow region 40 downstream of the shock waves 34 and 35 is characterized by severe turbulence and a large, non-symmetrical, time-varying wake vortex region 19. The flow velocity required to induce shock wave formation and the resulting complete boundary layer separation 40 for a given body 10 is dependent on the Reynolds number and the surface curvature.
There is strong coupling between the shock waves 34 and 35 and the separated flow 40. The turbulence and instabilities in the separated flow 40 are reflected in oscillatory instabilities in the strength and positions of the shock waves 34 and 35. Oscillatory motion of the shock wave 34 and 35 produces further turbulence in the separated flow 40 and wake vortex 19. The shock waves 34 and 35 on opposite sides of the body are also coupled through the separated flow and wake vortex. As the shock wave oscillations and flow turbulence levels increase, the two shock waves 34 and 35 begin to oscillate forward and backward in a coupled manner. When the bottom shock wave 35 increases in strength it moves backwards which increases the shock wave induced separated flow thickness 40 and pulls the wake vortex 19 down and off-center as illustrated in FIG. 1C. This causes the upper slightly weaker shock wave 34 to move backwards and the lower shock wave 35 to move forward until a condition such as that illustrated in FIG. 1D exists. The shock waves 34 and 35 will tend to oscillate back and forth between the two conditions illustrated in FIGS. 1C and 1D as indicated by the two-headed arrow 44.
The periodic and non-symmetric instability of the shock wave/boundary layer interaction results in large pressure fluctuations on the body surface 24. The pressure fluctuations produce a time-varying differential lift force and, in turn, impart a strong time-varying vibration or buffeting of the body 10. This vibration, directly related to the shock wave instabilities, can be very severe. Under some combinations of speed and body shape, the vibrations can result in structural failure leading to the potential loss of the aircraft and crew. In addition, the critical velocity M.sub.crit is highly dependent on the body shape, curvature, and angle of attack so practical aircraft may have several different critical velocities where vibration is a concern.
Much attention has been paid to buffeting in the transonic velocity range due to the potentially disastrous consequences to aircraft, crew and passengers. The approach to minimizing the buffeting has usually been concentrated on careful aerodynamic shaping of aircraft structures to minimize the existence of shock wave induced separation 40 and/or the avoidance of operating in the transonic region through the implementation of flight speed restrictions. However, it is usually impossible to produce airframe components that are always shape-optimized. Local supersonic conditions can occur at a number of subsonic aircraft speeds on most practical aircraft. This problem is exacerbated when pods or other devices are attached to the airframe as is often done for military purposes. Pods are often relatively thick structures falling into the class of bodies known in the art as bluff bodies and have freestream M.sub.crit values in the range of 0.4 to 0.7 Mach, much lower velocities than that of the aircraft alone. Vibrations in the pod structures can be so severe at particular velocities that the vibrations can result in the destruction of the pod to wing mounting structure and potentially the loss of the entire aircraft.
Methods for controlling unstable transonic flow and the induced vibrations have been implemented with varying success as an alternative to flight speed restrictions. Altering of airfoil and fuselage shapes has proven relatively successful for many airframe related problems. However, the shape of the pod is often dictated by other than aerodynamic considerations. In cases where non-aerodynamic considerations dictate shapes or in cases when shape optimization may be otherwise impractical, other techniques are needed.
One technique, known as Laminar Flow Control (LFC) discussed by Bertin and Smith, Aerodynamics for Engineers, Second Edition, Prentice Hall, 1989, pages 548-549, seeks to control turbulence and thereby control induced vibrations by maintaining the laminar flow condition. In LFC, laminar flow is maintained by the removal of the inner most part of the boundary layer using suction. LFC is normally used for drag reduction rather than transonic flow instability problems but can help delay the onset of severe turbulence.
Another method of preventing or delaying the onset of vibrations associated with the unstable shock waves is the use of vortex generators. This method seeks to prevent boundary layer separation by purposely inducing a shallow turbulent boundary layer (Bertin and Smith, 1989, pages 138-139) over the body surface. Vortex generators are rows of vertical fences arranged in an inclined manner relative to the flow that disrupt the laminar flow and produce a weakly turbulent boundary layer. The vortex generators have the effect of delaying the separation of flow associated with shock wave formation. Both of these methods attempt to prevent or delay the formation of the shock wave instead of addressing the instabilities in the shock wave that are the ultimate cause of the vibration problem.
It would be desirable to have an apparatus and/or method to stabilize the unstable shock waves 34 and 35 thereby preventing or minimizing induced vibrations to the body 10. A device or method that can be retrofitted or applied to aircraft structures such as those that typically experience vibrations during flight, would solve a long standing unsolved problem associated with transonic flight and known in the aerodynamic art.